The present invention relates generally to thermal management systems, and more specifically to a method of fabricating a structural panel with an integrated pumped-fluid loop thermal management system while maintaining the panel's stiffness-to-mass ratio.
Due to the vacuum of the space environment, traditional spacecraft thermal control techniques usually rely upon conduction and radiation to dissipate and reject heat generated by on-board electronic equipment. The spacecraft is designed such that a conduction path exists between each electronic component, the spacecraft bus, and a radiator outside the spacecraft, which radiates the electronics' waste heat to space. Unfortunately, the amount of heat which can be dissipated and rejected using this approach is severely limited, and it is difficult to efficiently transfer this heat over large distances. To maximize the effectiveness of this approach, electronic components are often mounted on the inside of a spacecraft structural panel. The outside of the panel serves as a radiator to space. Heat is conducted from the electronic components, through the panel, and to the radiator where it is rejected to space. This approach minimizes the distance over which the heat must be transported but severely restricts the placement of the electronic components within the spacecraft. Additionally, the amount of heat that can be removed before the component overheats is limited due to the relatively high thermal resistance of the structural panel. Several other technologies have been developed to help overcome these limitations. For example, heat pipes, loop heat pipes, and capillary-pumped loops are two-phase heat transfer devices that can transport significantly more heat a farther distance than most solid materials which rely on conduction alone. However, each of these devices adds a significant amount of weight and volume to the system. They also tend to be complicated and expensive and need to be custom designed for each spacecraft.
Additionally, the power levels of electronic components aboard spacecraft have risen dramatically over the years and will continue to do so in the future, while at the same time, spacecraft are becoming smaller and more compact. The result is much higher heat flux densities that must be dissipated by the thermal control system. These high densities can sometimes be mitigated using a thermal doubler to help spread the heat over a larger area, but doublers are not always sufficient and in some cases traditional techniques are inadequate to dissipate such fluxes. For example, state-of-the-art loop heat pipes are limited to heat flux capacities in the tens of watts per square centimeter, but many next-generation electronic components are expected to generate fluxes in the hundreds of watts per square centimeter. Other techniques, such as pumped fluid loops, may achieve considerably higher capacities, but to date these systems have added a significant amount of weight and complexity to the spacecraft thermal control system and have suffered from reliability issues.
Finally, the requirement to manage this increase in power and heat flux is compounded by the desire for modular, reconfigurable, and rapidly-deployable spacecraft. None of the aforementioned thermal management techniques meet these requirements, as each must be tailored to a specific application. The demands for higher power dissipation with increased heat flux capacity while being rapidly designed and integrated into a spacecraft bus are stretching the performance limits of traditional thermal management techniques. New technologies are required that can satisfy the thermal requirements of next-generation spacecraft without adding a significant amount of mass or volume to the thermal management system. One proposed technology that shows promise is addressed in U.S. patent application Ser. No. 12/049,474 filed Mar. 17, 2008 entitled, “Grid-stiffened Panel with Integrated Channels,” which discloses a grid-stiffened panel with fluid channels integrated in such a way as to preserve the original stiffness-to-mass ratio of the panel. A method for fabricating such a panel using composite materials is disclosed herein.